US9157456B2 - Method for monitoring the clearance of a kinematic link between a control member and a receiving member - Google Patents

Method for monitoring the clearance of a kinematic link between a control member and a receiving member Download PDF

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Publication number
US9157456B2
US9157456B2 US13/684,432 US201213684432A US9157456B2 US 9157456 B2 US9157456 B2 US 9157456B2 US 201213684432 A US201213684432 A US 201213684432A US 9157456 B2 US9157456 B2 US 9157456B2
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Prior art keywords
control member
blades
clearance
kinematic link
travel
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US13/684,432
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US20130136575A1 (en
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Nicolas Potel
Hichem Belghagi
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Safran Aircraft Engines SAS
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SNECMA SAS
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Assigned to SAFRAN AIRCRAFT ENGINES reassignment SAFRAN AIRCRAFT ENGINES CHANGE OF NAME (SEE DOCUMENT FOR DETAILS). Assignors: SNECMA
Assigned to SAFRAN AIRCRAFT ENGINES reassignment SAFRAN AIRCRAFT ENGINES CORRECTIVE ASSIGNMENT TO CORRECT THE COVER SHEET TO REMOVE APPLICATION NOS. 10250419, 10786507, 10786409, 12416418, 12531115, 12996294, 12094637 12416422 PREVIOUSLY RECORDED ON REEL 046479 FRAME 0807. ASSIGNOR(S) HEREBY CONFIRMS THE CHANGE OF NAME. Assignors: SNECMA
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D27/00—Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids
    • F04D27/002—Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids by varying geometry within the pumps, e.g. by adjusting vanes
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/54—Fluid-guiding means, e.g. diffusers
    • F04D29/56—Fluid-guiding means, e.g. diffusers adjustable
    • F04D29/563—Fluid-guiding means, e.g. diffusers adjustable specially adapted for elastic fluid pumps
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D17/00—Regulating or controlling by varying flow
    • F01D17/02—Arrangement of sensing elements
    • F01D17/04—Arrangement of sensing elements responsive to load
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D17/00—Regulating or controlling by varying flow
    • F01D17/02—Arrangement of sensing elements
    • F01D17/06—Arrangement of sensing elements responsive to speed
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D17/00—Regulating or controlling by varying flow
    • F01D17/02—Arrangement of sensing elements
    • F01D17/08—Arrangement of sensing elements responsive to condition of working-fluid, e.g. pressure
    • F01D17/085—Arrangement of sensing elements responsive to condition of working-fluid, e.g. pressure to temperature
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D17/00—Regulating or controlling by varying flow
    • F01D17/10—Final actuators
    • F01D17/12—Final actuators arranged in stator parts
    • F01D17/14—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits
    • F01D17/16—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
    • F01D17/162—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes for axial flow, i.e. the vanes turning around axes which are essentially perpendicular to the rotor centre line
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D21/00—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for
    • F01D21/003—Arrangements for testing or measuring
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
    • F02C9/16—Control of working fluid flow
    • F02C9/20—Control of working fluid flow by throttling; by adjusting vanes
    • F02C9/22—Control of working fluid flow by throttling; by adjusting vanes by adjusting turbine vanes
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
    • G01M15/00—Testing of engines
    • G01M15/14—Testing gas-turbine engines or jet-propulsion engines
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/80—Diagnostics

Definitions

  • the present invention relates to the field of turbomachines such as aircraft turbojet engines and, more particularly, to a method for monitoring the clearance of a kinematic link between a control member such as an actuating cylinder and one or more receiving members of the turbomachine, such as blades of the stator of a high pressure compressor which are able to be oriented.
  • the high pressure compressor comprises blade stages forming the stator of the compressor and which alternate with blade stages forming, in turn, the rotor of the compressor.
  • the blades of the stator although fixed in rotation relative to the longitudinal axis of the turbojet engine, have variable spacing or pitch so as to be able to be oriented relative to radial axes, perpendicular to the longitudinal axis of the turbojet engine.
  • a kinematic link is provided between a control member such as an actuator or linear actuating cylinder and the receiving members such as the stator blades to be oriented.
  • Said kinematic link which is essentially mechanical, is relatively complex since it consists of numerous parts which are articulated to one another to transmit the travel or path of the displacement in translation of the actuator into a pivoting of the blades about their respective axes.
  • FIGS. 1 and 2 One embodiment of such a known kinematic link is shown in FIGS. 1 and 2 and is described hereinafter, after the introduction to said figures.
  • the adjustment of the angular position of the blades of the stator is a function of the linear travel of the actuating cylinder such that a relation exists between said travel and the orientation of the blades, i.e. a specific (planar) position of the stator in the turbojet engine, a given pressure and a given acceleration/deceleration (and thus a rotational speed) correspond to a given angle of the blades.
  • Dead travel of the actuating cylinder is thus present before the actual orientation of the blades by the presence of the operational clearance of the kinematic system (including the different clearances between the parts). If said clearance is far too high, this may have serious consequences. More specifically, for such a compressor with stator blades with variable spacing, a clearance which is too high, for example as a result of wear of the sockets for retaining the blades, is not currently detected, nor is it able to be detected. In this manner, the stator blades may thus be offset at an angle and come into frictional contact with the rotor blades, which may result in a serious incident and destruction of the compressor.
  • said position oscillates about its set position with an oscillation amplitude which is less than the kinematic clearances and thus has no effect on the angular position of the blades. Said oscillation causes unnecessary wear of the different parts of the kinematic system in addition to the parts making up the actuating cylinder.
  • the monitoring of the clearance between the control member and the receiving members is carried out by performing a visual inspection of the state of the actuating cylinders and the kinematic link which is a lengthy and complicated operation requiring the stoppage of the engine, to assess the state of the “actuating cylinder-kinematic link-blade” assembly. More specifically, it is not easy to access the confined areas of the engine including said kinematic systems and requires a significant operation time. Moreover, an estimated measurement of the clearance is not determined during said inspections. If, during an inspection, it is judged that the clearance may lead to malfunctioning, then adjustment operations are carried out on the relevant parts with the replacement thereof if required.
  • the object of the present invention is to remedy said drawbacks and the invention relates to a method making it possible to monitor the clearance of the kinematic link without having recourse to a lengthy immobilization of the engine and to awkward servicing operations.
  • the method for monitoring the clearance of a kinematic link between a control member such as a linear actuator and receiving members of a turbomachine, such as the blades of a stator with variable spacing of a compressor is noteworthy according to the invention in that it consists, based on an operating parameter of the turbomachine, in:
  • the clearance is monitored by simple linear displacements of the control member, such as the actuator, respectively a forward displacement, taking up the clearance to act directly on the receiving members such as the blades, and a return displacement until the value of the selected operating parameter is varied, from the travel corresponding to the total clearance produced.
  • the control member such as the actuator
  • a forward displacement taking up the clearance to act directly on the receiving members such as the blades
  • a return displacement until the value of the selected operating parameter is varied, from the travel corresponding to the total clearance produced.
  • the clearance is considered as acceptable and as not requiring intervention.
  • the implementation of the method for monitoring the clearance is particularly rapid and reliable by a simple operation of the control member and the turbojet engine, without dismantling said engine and without recourse to additional tools, which reduces the maintenance costs. It is also observed that the monitoring method ultimately does not require a direct measurement of the clearance, the variation of the value of the parameter under consideration being sufficient to permit the comparison of the travel of the control member corresponding to the clearance of the kinematic link with a reference travel.
  • the rotational speed of the compressor with the variable blade assembly is selected as an operating parameter of the turbojet engine.
  • the return travel of the linear actuator to the second position is stopped and the value of the travel between the two positions corresponding substantially to the total clearance of the kinematic system is ascertained and then compared.
  • parameters in terms of operating parameters of the turbojet engine may be selected and may relate to the pressure of the air circulating in the compressor with the variable blade assembly or the temperature of the exhaust gases.
  • the steps of the method are repeated according to an established timescale corresponding to a specific number of flights of an aircraft provided with turbojet engines.
  • FIG. 1 shows a partial and schematic longitudinal section of a turbojet engine of an aircraft showing, in particular, the high pressure compressor with its rotor with rotating blades and its stator with fixed blades, with variable spacing, and the kinematic link controlling said blades.
  • FIG. 2 is a perspective view of the kinematic link controlling the blades with variable spacing of the stator.
  • FIG. 3 is a graph showing the rotation of the stator blades as a function of the displacement path of the engine member, including the operational clearance of the kinematic link.
  • FIGS. 4 , 5 and 6 are graphs showing the implementation of the method according to the invention, respectively according to the position of the engine member, the position of the receiving members and the rotational speed of the engine.
  • the turbojet engine 1 shown partially in FIG. 1 such as a double-flow turbojet engine for an aircraft, shows the high pressure compressor 2 comprising, in the usual manner, fixed to the casing 3 thereof, the stator stages 4 with the blades 5 thereof at variable spacing, and therebetween, the rotor stages 6 with the rotating blades 7 thereof.
  • the primary flow F 1 of the air entering the turbojet engine 1 passes through the successive and alternate stages of the rotor 6 and the stator 4 of the compressor, whilst the secondary flow F 2 passes through the turbojet engine between an intermediate casing 8 surrounding the compressor and an external casing, not visible in the figures, of the turbojet engine.
  • the stator 2 constitutes the fixed part that the primary flow of air passes through.
  • the blades 7 of the rotor 6 are connected together by platforms 9 and are arranged so as to rotate, during the operation of the turbojet engine 1 , between the blades 5 of the stator 4 which are fixed in rotation relative to the blades 6 of the rotor but may be rotated about their own axis A, i.e. radially relative to the axis of rotation of the blades of the rotor, corresponding to the longitudinal axis, not visible in FIG. 1 , of the turbojet engine 1 .
  • said blades 5 are controlled by a mechanical system 10 provided between the intermediate casing 8 and the casing 3 of the compressor to which the stator stages are associated.
  • the system 10 illustrated in FIGS. 2 and 3 comprises a control member 11 such as an actuator or linear actuating cylinder, and a kinematic link 12 connecting the control member 11 to the stator blades 5 which serve as receiving members.
  • a control member 11 such as an actuator or linear actuating cylinder
  • a kinematic link 12 connecting the control member 11 to the stator blades 5 which serve as receiving members.
  • Said kinematic link 12 comprises a plurality of parts connected together and in particular, starting from the actuating cylinder, a control lever 14 , a control shaft 15 , a cylindrical transmission element 16 and a transmission structure 17 onto which is mounted a plurality of actuating links 18 connected to respective control rings 19 of the blades 5 , each of the rings 19 being connected via an orientation lever 20 to the blades 5 of the stator 4 , of which the feet are received in connecting sockets 21 housed in the casing 3 .
  • the transmission structure 17 to the blades 5 of the stator 4 in turn comprises a plurality of parts such as, in particular, a base 22 in which links 23 are articulated for transmitting force to the actuating links 18 fixed to the control rings 19 , and a support plate 24 receiving the control shaft 15 by a ball joint to retain the control shaft in the plate 24 and permit the rotation of the shaft 15 therein.
  • control actuating cylinder 11 causes the displacement of the entire kinematic link 12 of the system 10 to the receiving members, i.e. the orientation of the blades 5 of the stator 4 about their axis A.
  • each blade 5 is connected to the ring 19 by the orientation lever 20 such that the rotation of the ring 19 causes the displacement of the orientation levers 20 in rotation, said rotation permitting the angular orientation of the blades 5 in the sockets 21 relative to the primary flow F 1 .
  • the graph of FIG. 3 shows the dead travel C of the rod of the actuating cylinder 11 before the rotation of the blades 5 is produced about the axes A, said dead travel to a certain extent taking up the various clearances of the parts making up the kinematic link 12 from the actuating cylinder 11 to the blades 5 .
  • the travel C of the actuating cylinder is seen on the abscissa of the graph, during which the rotation of the stator blades 5 , shown on the ordinate, is zero, no rotation being produced.
  • the displacement path of the actuating cylinder 11 continues and proportionally causes, along a straight line D, the angular orientation of the blades 5 according to the established law as specified above.
  • FIGS. 4 , 5 and 6 which method may be implemented when the aircraft is grounded, without recourse to specific related installations. More specifically, as FIG. 4 shows, it is necessary to displace the rod of the linear actuating cylinder 11 (control member) from any initial position P which the actuating cylinder occupies to a first position P 1 where it is certain that the blades 5 of variable spacing of the stator 4 have been rotated, i.e. a rotation thereof about their radial axis A has indeed taken place. The clearance of the kinematic system 12 which is thus present between the initial position P and the first position P 1 has thus been taken up, since the rotation of the blades 5 is achieved whatever the angular variation implemented, said angular variation simply having to be effective.
  • the estimated measurement of the travel of the actuating cylinder between its two positions P 1 and P 2 is deduced from this displacement to and fro of the actuating cylinder 11 , therefore, namely a measurement which corresponds to the total clearance J of the kinematic link 12 by taking the rotational speed of the compressor as the operating parameter of the turbojet engine 1 , in this example for implementing the method.
  • the method consists in comparing the value of this ascertained measurement with a limit reference value previously determined by calculation and, if said ascertained measured value is beyond the theoretically admissible limit value, then a visual inspection of the kinematic link 12 is carried out and if required, an intervention with adjustments and potentially the replacement of the relevant parts of the transmission.
  • the measured value is less than the limit value, no inspection or intervention is required, meaning that the kinematic link is operational. It is possible to observe the ease with which said kinematic system may be monitored without recourse to dismantling the turbojet engine, with the resulting loss of time, or any direct measurement of the clearance in the region of said parts and thus any specific equipment, resulting in a saving of time and costs.
  • such a method for monitoring may be programmed regularly, for preventative purposes, after a certain number of flights of the aircraft, so as to be able to monitor said kinematic link ultimately over the entire life of the aircraft.
  • the clearance of the kinematic link is monitored between the actuating cylinder and the blades based on the rotational speed of the compressor in order to deduce therefrom the desired clearance, by measuring the difference in the position of the actuating cylinder P 1 and P 2 , and then to compare said clearance with a limit value.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • General Physics & Mathematics (AREA)
  • Geometry (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Length Measuring Devices With Unspecified Measuring Means (AREA)
US13/684,432 2011-11-24 2012-11-23 Method for monitoring the clearance of a kinematic link between a control member and a receiving member Active 2034-04-22 US9157456B2 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR1160771A FR2983236B1 (fr) 2011-11-24 2011-11-24 Procede de surveillance du jeu d'une cinematique de liaison entre un organe de commande et un organe recepteur.
FR1160771 2011-11-24

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US20130136575A1 US20130136575A1 (en) 2013-05-30
US9157456B2 true US9157456B2 (en) 2015-10-13

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FR (1) FR2983236B1 (fr)
GB (1) GB2496984B (fr)

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB201405572D0 (en) * 2014-03-28 2014-05-14 Rolls Royce Plc Actuation system investigation apparatus
CN105570196B (zh) * 2014-10-31 2019-09-06 特灵国际有限公司 致动进口导叶的连杆机构
FR3094790B1 (fr) * 2019-04-05 2021-03-05 Safran Aircraft Engines Système de surveillance d’un système d’actionnement d’un élément d’une turbomachine
US11560810B1 (en) * 2021-07-20 2023-01-24 Rolls-Royce North American Technologies Inc. Variable vane actuation system and method for gas turbine engine performance management

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US20090123272A1 (en) 2007-11-13 2009-05-14 Love Andrew C Adaptive variable geometry turbocharger strategy
US7689376B2 (en) * 2008-05-29 2010-03-30 Honeywell International Inc Method of calibrating an actuator system for a variable nozzle of a turbocharger
US7855525B2 (en) * 2007-10-30 2010-12-21 Delphi Technologies, Inc. Method for controlling a holding force against, and limiting impact with travel limit positions
FR2947310A1 (fr) 2009-06-26 2010-12-31 Snecma Dispositif et methode de positionnement d'un equipement a geometrie variable pour une turbomachine, utilisant un verin a mesure relative.
EP2383439A2 (fr) 2010-04-28 2011-11-02 General Electric Company Systèmes, procédés et appareil permettant de contrôler les positions d'aube de guidage de turbine
US8690521B2 (en) * 2008-09-30 2014-04-08 Snecma System for controlling variable geometry equipment for a turbine engine, especially by bellcranks

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US6341238B1 (en) * 1998-10-01 2002-01-22 United Technologies Corporation Robust engine variable vane monitor logic

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US7855525B2 (en) * 2007-10-30 2010-12-21 Delphi Technologies, Inc. Method for controlling a holding force against, and limiting impact with travel limit positions
US20090123272A1 (en) 2007-11-13 2009-05-14 Love Andrew C Adaptive variable geometry turbocharger strategy
EP2067957A2 (fr) 2007-11-13 2009-06-10 Honeywell International Inc. Stratégie adaptative pour turbocompresseur à géométrie variable
US7689376B2 (en) * 2008-05-29 2010-03-30 Honeywell International Inc Method of calibrating an actuator system for a variable nozzle of a turbocharger
US8690521B2 (en) * 2008-09-30 2014-04-08 Snecma System for controlling variable geometry equipment for a turbine engine, especially by bellcranks
FR2947310A1 (fr) 2009-06-26 2010-12-31 Snecma Dispositif et methode de positionnement d'un equipement a geometrie variable pour une turbomachine, utilisant un verin a mesure relative.
US20120107088A1 (en) 2009-06-26 2012-05-03 Snecma Device and method for positioning variable-geometry equipment for a turbomachine, using a relative-measurement jack
EP2383439A2 (fr) 2010-04-28 2011-11-02 General Electric Company Systèmes, procédés et appareil permettant de contrôler les positions d'aube de guidage de turbine
US20110268554A1 (en) 2010-04-28 2011-11-03 General Electric Company Systems, methods, and apparatus for controlling turbine guide vane positions
US8770912B2 (en) * 2010-04-28 2014-07-08 General Electric Company Systems, methods, and apparatus for controlling turbine guide vane positions

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Also Published As

Publication number Publication date
GB201221067D0 (en) 2013-01-09
FR2983236B1 (fr) 2016-05-06
FR2983236A1 (fr) 2013-05-31
GB2496984B (en) 2017-10-04
GB2496984A (en) 2013-05-29
US20130136575A1 (en) 2013-05-30

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